The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in high and low pressure turbines which power the compressor and a fan in an exemplary turbofan aircraft engine configuration.
The first stage turbine rotor blades receive hot combustion gases from their corresponding turbine nozzle directly from the combustor. The turbine blades are therefore made of superalloy metal having enhanced strength at the high operating temperature of the turbine. And, the blades are typically hollow and internally cooled by bleeding a portion of pressurized air from the compressor.
The cooling air bled from the compressor bypasses the combustor and correspondingly reduces efficiency of the engine. Accordingly, the total amount of air bled from the compressor for cooling the turbine blades should be minimized while correspondingly maximizing the useful life thereof which can reach thousands of hours of operation, with corresponding periodic inspection and maintenance.
The prior art is replete with many patents having myriad configurations for efficiently cooling turbine rotor blades which are based on the particular configuration of the engine and operating cycle. The complexity of blade cooling configurations is also based on the varying distribution of pressure and heat loads around the concave pressure side of the airfoil and the convex suction side which extend radially in span from root to tip and axially in chord between opposite leading and trailing edges.
The life of the typical turbine blade is limited by any local hot spot experienced thereby which increases the thermal distress experienced by the blade eventually leading to undesirable airfoil cracking for example. The airfoil has relatively thin walls, and any thermal crack developed therein during operation can lead to undesirable leakage of the internal cooling air.
Airfoil cracking due to thermal distress typically occurs at the airfoil tip which is subject to heating by the combustion gases not only from the opposite pressure and suction sides of the airfoil, but also from combustion gases leaking over the tip in the small radial clearance or gap provided with the surrounding turbine shroud.
Accordingly, it is desired to provide a turbine blade having an improved cooling configuration for accommodating tip cracking experienced in later blade life.